The aeronautical industry requires structures which, on the one hand, support the loads to which they are subjected fulfilling high stiffness and resistance demands and, on the other hand, are as light as possible. A consequence of this requirement is the continuously expanding use of composite materials in primary structures because, by conveniently applying these materials, an important weight reduction can be achieved compared with structures designed with metallic materials.
The composite materials that are most used in the aeronautical industry consist of fibers or fiber bundles embedded in a matrix of thermosetting or thermoplastic resin, in the form of a preimpregnated or “prepreg” material. Its main advantages refer to:                Their high specific strength with respect to metallic materials. It is the strength/weight equation.        Their excellent behaviour under fatigue loads.        The possibilities of structural optimization thanks to the anisotropy of the material and the possibility of combining fibers with different orientations, allowing the design of the elements with different mechanical properties adjusted to the different needs in terms of applied loads.        
As is well known, the main structural elements of aircraft fuselages are the skin, the frames and the stringers. The skin is stiffened longitudinally with stringers to reduce the skin thickness, making it more competitive in terms of weight, while the frames avoid the overall instability of the fuselage and can be subjected to the introduction of local loads. Other structural elements can be found inside an aircraft fuselage, such as beams, which act as a frame for open sections of the fuselage or which are used to withstand the loads introduced by the cabin floor of the aircraft.
The fuselage structure made out of composite materials which is nowadays more commonly used consists, on the one hand, of a skin with integrated stringers, co-bonded or co-cured, and on the other hand, of complete or floating frames which are manufactured separately and which are then riveted to the fuselage skin. The document U.S. Pat. No. 5,242,523 describes a structure such as this one combining the use of omega-shaped stringers with C-shaped frames.
Omega-shaped stringers have been widely used in fuselages in the past few years because they have a high inertia and can provide support and stability to a great skin panel due to its geometry. These characteristics, along with the advantages it presents for its manufacture, due to the simplification and reduction of the tooling, and therefore of cost, make its use in the reinforcement of fuselage skins very interesting.
Regarding frames both open and closed section frames have been proposed for fuselages. Open section shaped frames such as the frames used in U.S. Pat. No. 5,242,523 are indeed convenient from a manufacturing point of view although require high stabilized webs while closed section shaped frames, such omega-shaped frames, that have a higher strength, raise manufacturing problems.
The present invention focuses on finding a solution for these drawbacks.